In an aircraft gas turbine engine, air is drawn into the front of the engine, compressed by a shaft-mounted compressor, and mixed with fuel. The mixture is burned, and the hot exhaust gases are passed through a turbine mounted on the same shaft. The flow of combustion gas turns the turbine by impingement against the airfoil section of the turbine blades, which turns the shaft and provides power to the compressor. The hot exhaust gases flow from the back of the engine, driving it and the aircraft forward. The hotter the combustion and exhaust gases, the more efficient is the operation of the jet engine. Thus, there is incentive to raise the combustion gas temperature.
The compressors and turbine of the turbine engine can comprise turbine disks (sometimes termed “turbine rotors”) or turbine shafts, as well as a number of blades mounted to the turbine disks/shafts and extending radially outwardly therefrom into the gas flow path, and rotating. Also included in the turbine engine are rotating, as well as static, seal elements that channel the airflow used for cooling certain components such as turbine blades and vanes. The airflow channeled by these rotating, as well as static, seal elements carry corrodant deposits to the non-gas path sides of turbine blades. As the maximum operating temperature of the turbine engine increases, the turbine blades are subjected to higher temperatures. As a result, oxidation and corrosion of the turbine blades have become of greater concern.
Metal salts such as alkaline sulfate, sulfites, chlorides, carbonates, oxides, and other corrodant salt deposits resulting from ingested dirt, fly ash, volcanic ash, concrete dust, sand, sea salt, etc. are a major source of the corrosion, but other elements in the bleed gas environment can also accelerate the corrosion. Alkaline sulfate corrosion in the temperature range and atmospheric region of interest results in pitting of the turbine blade substrate at temperatures typically starting around 1200° F. (649° C.). This pitting corrosion has been shown to occur on turbine blades, primarily the region beneath platforms of turbine blades. The oxidation and corrosion damage can lead to failure or premature removal and replacement of the turbine blades unless the damage is reduced or repaired.
Turbine blades for use at the highest operating temperatures are typically made of nickel-base superalloys selected for good elevated temperature toughness and fatigue resistance. In addition, the turbine blade alloys are coated with environmental coatings to primarily protect the turbine airfoil and platform structures for oxidation and corrosion. These coatings may additionally be deposited on the under platform region of the turbine blade. Typical environmental coatings in wide use include MCrAlX overlay coatings (where M is iron, cobalt and/or nickel, and X is yttrium or another rare earth element), and diffusion coatings that contain aluminum intermetallics, predominantly beta-phase nickel aluminide (βNiAl) and platinum aluminides (PtAl). These superalloys and the existing environmental coatings used have resistance to oxidation and corrosion damage, but that resistance is not sufficient to protect them at sustained operating temperatures now being reached in gas turbine engines.
Corrosion resistant coating compositions have been suggested for use with various gas turbine components. These include aqueous corrosion resistant coating compositions comprising phosphate/chromate binder systems and aluminum/alumina particles. See, for example, U.S. Pat. No. 4,606,967 (Mosser), issued Aug. 19, 1986 (spheroidal aluminum particles); and U.S. Pat. No. 4,544,408 (Mosser et al), issued Oct. 1, 1985 (dispersible hydrated alumina particles). Corrosion resistant diffusion coatings can also be formed from chromium, or from the respective oxide (i.e., chromia). See, for example, commonly assigned U.S. Pat. No. 6,283,715 (Nagaraj et al), issued Sep. 4, 2001 (chromium diffusion coating). A number of corrosion resistant coatings have also been specifically considered for use on turbine disk/shaft and seal elements. See, for example, U.S. Patent Application 2004/0013802 A1 (Ackerman et al), published Jan. 22, 2004 (metal-organic chemical vapor deposition of aluminum, silicon, tantalum, titanium or chromium oxide on turbine disks and seal elements to provide a protective coating). These prior corrosion resistant coatings have a number of disadvantages when used with turbine blades, including: (1) possibly adversely affecting the fatigue life of the turbine blade elements, especially when these prior coatings diffuse into the underlying metal substrate; (2) potential coefficient of thermal expansion (hereinafter, “CTE”) mismatches between the coating and the underlying metal substrate that can make the coating more prone to spalling; and (3) more complicated and expensive processes (e.g., chemical vapor deposition) for applying the corrosion resistant coating to the metal substrate.
What is needed are coatings and coating compositions for turbine blades that: (1) provide corrosion resistance, especially at higher or elevated temperatures; (2) do not affect other mechanical properties of the underlying metal substrate or potentially causing other undesired effects such as spalling; (3) can be formed by relatively uncomplicated and inexpensive methods; (4) can allow for non-destructive evaluation of the underlying substrate during engine overhaul; and (5) can be reapplied or refurbished for continued engine operation. The present invention provides these and other related advantages.